The present invention relates to axial flow compressors for use in gas turbines or industrial applications, and in particular, it relates to axial flow compressors having axial compressor blades that exhibit high performance and low pressure loss.
Conventional axial flow compressors have heretofore employed NACA-65 profiles which have been developed for use in subsonic applications as has been described in the document "NASA, SP-36 (AERODYNAMIC DESIGN OF AXIAL-FLOW COMPRESSORS )"1965. In recent years, however, demands for higher pressure rate and improved efficiency are necessitating various attempts to increase the velocity at the inlet of the blade rows.
Regarding transonic blade rows, a double circular arc profile in which the suction surface and pressure surface are composed of single circular arcs, respectively, is described in the literature "Pumps and Blowers: Theories and Applications", 343rd Conference (1971), Japan Society of Mechanical Engineers.
The conventional axial compressors have a problem in that pressure loss due to shock waves occurring on the blade surface becomes excessively great when the inlet Mach number exceeds the critical Math number, thereby lowering efficiency substantially. Further, blade profile losses tend to increase as a result of the shock wave arising with an increase in the velocity at the inlet of the blade rows.
Therefore, it is important to provide for blade profiles which exhibit high performance, in particularly, when the inlet flow is in a higher subsonic region and in a transonic range.